The present invention relates generally to turbine blades for a gas turbine engine and, in particular, to the cooling of the tip and the tip leakage flow of such turbine blades.
It is well known that air is pressurized in a compressor of a gas turbine engine and mixed with fuel in a combustor to generate hot combustion gases, whereupon such gases flow downstream through one or more turbines so that energy can be extracted therefrom. In accordance with such turbine, a row of circumferentially spaced apart rotor blades extend radially outwardly from a supporting rotor disk. Each blade typically includes a dovetail which permits assembly and disassembly of the blade in a corresponding dovetail slot in the rotor disk, as well as an airfoil which extends radially outwardly from the dovetail.
The airfoil has a generally concave pressure side and generally convex suction side extending axially between corresponding leading and trailing edges and radially between a root and a tip. It will be understood that the blade tip is spaced closely to a radially outer turbine shroud for minimizing leakage therebetween of the combustion gases flowing downstream between the turbine blades. Maximum efficiency of the engine is obtained by minimizing the tip clearance or gap, but is limited by the differential thermal and mechanical expansion and contraction between the rotor blades and the turbine shroud for reducing the likelihood of undesirable tip rubs.
Since the turbine blades are bathed in hot combustion gases, effective cooling is required for ensuring a useful life. The blade airfoils are hollow and disposed in flow communication with the compressor so that a portion of pressurized air bled therefrom is received for use in cooling the airfoils. Airfoil cooling is quite sophisticated and may be effected using various forms of internal cooling channels and features, as well as cooling holes through the walls of the airfoil for discharging the cooling air.
The airfoil tip is particularly difficult to cool since it is located directly adjacent to the turbine shroud and the hot combustion gases which flow through the tip gap therebetween. Accordingly, a portion of the air channeled inside the airfoil is typically discharged through the tip for cooling thereof. The tip typically includes a continuous radially outwardly projecting edge rib disposed coextensively along the pressure and suction sides between the leading and trailing edges, where the tip rib follows the aerodynamic contour around the airfoil and is a significant contributor to the aerodynamic efficiency thereof.
Generally, the tip rib has portions spaced apart on the opposite pressure and suction sides to define an open top tip cavity. A tip plate or floor extends between the pressure and suction side ribs and encloses the top of the airfoil for containing the cooling air therein. Tip holes are also provided which extend through the floor for cooling the tip and filling the tip cavity.
It will be appreciated that several exemplary patents related to the cooling of turbine blade tips are disclosed in the art, including: U.S. Pat. No. 5,261,789 to Butts et al.; U.S. Pat. No. 6,179,556 to Bunker; U.S. Pat. No. 6,190,129 to Mayer et al.; and, U.S. Pat. No. 6,059,530 to Lee. These patents disclose various blade tip configurations which include an offset on the pressure and/or suction sides in order to increase flow resistance through the tip gap. Nevertheless, improvement in the pressure distribution near the tip region is still sought to further reduce the overall tip leakage flow and thereby increase turbine efficiency.
Thus, in light of the foregoing, it would be desirable for a turbine blade tip to be developed which alters the pressure distribution near the tip region to reduce the overall tip leakage flow and thereby increase the efficiency of the turbine. It is also desirable for such turbine blade tip to develop one or more recirculation zones adjacent the ribs at such tip in order to improve the flow characteristics and pressure distribution at the tip region.
In a first exemplary embodiment of the invention, a turbine blade for a gas turbine engine is disclosed as including an airfoil and integral dovetail for mounting the airfoil along a radial axis to a rotor disk inboard of a turbine shroud. The airfoil further includes: first and second sidewalls joined together at a leading edge and a trailing edge, where the first and second sidewalls extend from a root disposed adjacent the dovetail to a tip plate for channeling combustion gases thereover; and, at least one tip rib extending outwardly from the tip plate between the leading and trailing edges. The tip rib is oriented so that an axis extending longitudinally therethrough is at an angle with respect to the radial axis for at least a designated portion of an axial length of the turbine blade. The angle between the longitudinal axis and the radial axis may be substantially the same across the designated portion or may vary thereacross.
In a second exemplary embodiment of the invention, a turbine blade for a gas turbine engine is disclosed as including an airfoil and integral dovetail for mounting the airfoil along a radial axis to a rotor disk inboard of a turbine shroud. The airfoil further includes: first and second sidewalls joined together at a leading edge and a trailing edge, where the first and second sidewalls extend from a root disposed adjacent the dovetail to a tip plate for channeling combustion gases thereover; and, at least one tip rib extending outwardly from the tip plate between the leading and trailing edges. The tip rib is oriented with respect to the radial axis so that a first recirculation zone of the combustion gases is formed adjacent a distal end of the tip rib which reduces a leakage flow of the combustion gases between the airfoil and the shroud for at least a designated portion of an axial length of the turbine blade.